This invention relates to control of pitch and yaw of vehicles propelled by high-velocity gas jets, as for example rockets.
Rocket and other jet-type propulsion systems are widely used for vehicle propulsion, notably for lifting payloads into orbit, for destructive military missiles, and for military aircraft. In general, such systems or engines produce thrust by discharge of a plume or exhaust at high velocity along the axis of a nozzle. The problems of pitch and yaw directional control of such vehicles are widely known, and have been solved in a number of ways. External fins and canards have been used for directional control. These fins may be fixed for the most general directional control, or they may include articulable flaps which are controlled in response to a controller for feedback flight control. When the vehicle equipped with fins or canards must itself be carried on an aircraft before launch, or mounted in a canister, the fins andor canards may be arranged in a stowed configuration, and deployed in conjunction with the initial stages of launch. Such fins and canards depend upon aerodynamic forces, so are only usable within the atmosphere, which may be taken to extend to an altitude of 100,000 feet. Within the atmosphere, the use of fins or canards increases the drag of the vehicle equipped therewith.
Another form of attitude or directional control of a vehicle equipped with rocket-type propulsion is that of thrust vector control (TVC), described U.S. Pat. No. 2,943,821, issued Jul. 5, 1960 in the name of Wetherbee, Jr.; U.S. Pat. No. 3,166,897, issued Aug. 21, 1961 in the name of Lawrence et al; in U.S. Pat. No. 3,132,476, issued May 12, 1964 in the name of Conrad; and U.S. Pat. No. 3,132,478, issued May 12, 1964 in the name of Thielman, and in the text xe2x80x9cRocket Propulsion Elementsxe2x80x9d by Robert Sutton. Sutton categorizes TVC mechanisms into four basic categories:
(a) Mechanical deflection of a nozzle or thrust chamber;
(b) insertion or adjustment of vanes located in the jet exhaust stream;
(c) injection of fluid into the diverging nozzle section to deflect the exhaust flow; and
(d) separate thrust-producing devices which are independent of the main flow through the nozzle, providing two thrust vectors which may be summed to obtain the net thrust vector.
It should be noted that this last may not be a form of TVC, since it does not act on the thrust vector itself, but merely adds a separate thrust vector.
Mechanical deflection of a nozzle or thrust chamber requires a highly reliable movable structure which is subject to the entire thrust load, which may be costly and undesirably massive. Insertion or adjustment of vanes within the exhaust stream requires vanes which are structurally sound at the very high temperatures and pressures of the exhaust stream. Thrust-producing devices independent of the main nozzle have been used, especially for end-of-flight corrections of destructive missiles acting against maneuvering targets; they must, however, be located within the body of the vehicle if additional aerodynamic drag is to be avoided.
According to Sutton, xe2x80x9cthe injection of secondary fluid through the wall of the nozzle into the main gas stream has the effect of forming oblique shocks in the nozzle diverging section, thus causing a deflection of part of a main gas flow,xe2x80x9d and this deflection of the main gas flow, in turn, results in a deviation of the thrust vector from the axis of the nozzle.
Liquid injection thrust vector control is described in U.S. Pat. No. 3,737,103, issued Jun. 5, 1973 in the name of Howell et al. Liquid injection thrust vector control is a proven technology, which is used in applications such as Titan III and Minuteman. In liquid injection thrust vector control, liquid is stored in either the propellant tanks or auxiliary tanks of the vehicle. The liquid is controllably distributed or manifolded to various injection positions around the periphery of the nozzle. When a pitch or yaw correction is desired, a signal is sent to a valve or valves controlling the injection of liquid into the exhaust plume at locations associated with the plane(s) of the correction thrust. Injection of the liquid into the exhaust stream results in vaporization of the liquid, and also results in a change in thrust along the relevant plane. Liquid injection has known problems, which include the instability of stored liquids, as described in U.S. Pat. No. 3,092,963, issued Jun. 11, 1963 in the name of Lawrence. Also, the axial thrust of the vehicle is reduced by the energy required to vaporize the injected liquid, and to bring it up to the temperature of the surrounding gas. The amount of liquid which is required to produce a given change in attitude is generally determined by experimentation.
FIG. 1 is a chart illustrating the amount of side injectant which is required to produce a side force, according to Sutton. In FIG. 1, the ordinate- or y-axis represents the ratio of the side force divided by (or normalized to) axial force, and the abscissa- or x-axis represents the ratio of injectant mass flow divided by primary mass flow. As illustrated in FIG. 1, injection of inert liquids results in the least side thrust or directional control for a given mass flow, while reactive fluids provide greater control. The greatest control is provided by a flow of propellant hot gas. Such charts can be obtained experimentally by maintaining a constant main exhaust flow rate through the nozzle, while varying the side injection flow rate.
As suggested by FIG. 1, greater side force or thrust control can be achieved by injection of reactive liquids than of inert liquids. U.S. Pat. No. 2,952,123, issued Sep. 3, 1960 in the name of Rich, describes injection of fuel into a jet nozzle, which burns in the supersonic exhaust stream to provide directional control. Similarly, FIG. 1 indicates that injection of propellant hot gas provides yet greater side force or directional control as a function of mass flow.
As an example of the use of the chart of FIG. 1, consider a rocket engine or motor which generates an axial thrust of 10,000 pounds force (lbf) at an exhaust flow rate of 33 pounds of mass (lbm). If a 2xc2x0 deflection of the thrust vector is required, then (10,000) (sin 2xc2x0) of lateral force, corresponding to 349 lbf of normal side force, is required. The side-to-axial-force ratio is calculated as 0.035, which corresponds to a side injection flow rate of (0.035xc3x9733)=1.98 lbm/s for inert liquids, 1.32 lbm/s for reactive liquids, or 1.0 lbm/s for propellant hot gas.
From FIG. 1, it is apparent that injection of hot gas is the most effective way, in terms of relative mass flow, to achieve side force or directional control. One advantageous way to provide hot gas for side injection is to tap the gas from the main combustion chamber, because the chamber pressure is greater than the static pressure in the nozzle as a result of expansion, and a substantial side injection flow rate can therefore be achieved. U.S. Pat. No. 3,759,039, issued Sep. 18, 1973 in the name of Williams, describes the bleeding of hot gases from the combustion chamber of a rocket, by way of controllable valves, into the side of the nozzle. In such an arrangement, the valves must control the flow of very hot gases, which may adversely affect their reliability, and may result in a costly structure using exotic materials and sacrificial elements or coatings.
The problem of control of the flow of very hot propellant-type gases makes the use of cooled or cold gas advantageous. The chamber temperatures of liquid- and solid-propellant rockets may approach 6000xc2x0 F., which is too high for conventional piping and valves. Cool-gas or cooled-gas injection is described in U.S. Pat. No. 3,255,971, issued Jun. 14, 1966 in the name of Widell; U.S. Pat. No. 3,698,642, issued Oct. 17, 1972, in the name of McCullough; U.S. Pat. No. 4,384,694 in the name of Watanabe et al.; U.S. Pat. No. 4,424,670, issued Jan. 10, 1984 in the name of Calabro; and in the abovementioned Lawrence patent. The hot gas flowing through the valves in the cool- or cooled-gas injection arrangements should be at a temperature no greater than about 1100xc2x0 F. In liquid-propellant rocket engines, attempts have been made to draw the hot gases from the fuel-rich boundary layer at the chamber wall, which is often at a lower temperature than the average chamber temperature. In solid-propellant rockets, less aggressive propellants, containing fewer oxidizing ingredients, can be used, but at the expense of reducing already-limited performance.
Improved thrust vector control is desired.
A vehicle according to an aspect of the invention comprises a source, such as a rocket engine chamber, of main propulsion fluid, and a nozzle coupled to the source of propulsion fluid, for generating propulsion thrust by discharge of the main propulsion fluid generally along a discharge axis. A hybrid fluid generator includes a solid xe2x80x9cpropellantxe2x80x9d or fuel grain, which can be combusted in the presence of oxidizer. The hybrid exhaust fluid generator includes an exit port coupled to a side of the nozzle at a first location, and also including an oxidizer input port. The hybrid exhaust fluid generator generates secondary fluid at the exit port in response to reaction of the grain with oxidizer applied to the oxidizer input port, and injects the secondary fluid into the side of the nozzle, for thereby deflecting the main propulsion thrust relative to the axis. In this arrangement, the amount of the deflection is controlled by control of the rate of flow of the oxidizer to the oxidizer input port of the hybrid exhaust fluid generator.
Another version of a vehicle according to the invention comprises a further or second hybrid exhaust fluid generator. The further hybrid exhaust fluid generator is similar to the first, in that it includes a solid grain and an exit port coupled to the side of the nozzle. In the case of the second hybrid exhaust fluid generator, the exit port is coupled to the nozzle at a position spaced about, and in one embodiment diametrically opposite, relative to the axis, to the first location. The second hybrid exhaust fluid generator also including an oxidizer input port, and is for generating further secondary fluid at the exit port of the further hybrid exhaust fluid generator in response to reaction of the oxidizer with the grain, and for injecting the further secondary fluid into the side of the nozzle at the diametrically opposite location, for thereby deflecting the main propulsion thrust, relative to the axis, in a direction opposed to that of the first-mentioned hybrid exhaust fluid generator. A particular manifestation includes control arrangement coupled to the oxidizer input ports of the first-mentioned and further hybrid exhaust fluid generators, for, when thrust deviation is desired within a plane including the axis and the first-mentioned hybrid exhaust fluid generator, providing one of the first-mentioned and further hybrid exhaust fluid generators with oxidizer. Those skilled in the art know that this provides thrust deflection in the plane of the two exit ports, depending upon control of the secondary fluid flow rates. In a preferred embodiment, the control arrangement couples the oxidizer to the one of the first-mentioned and further hybrid exhaust fluid generators to the exclusion of the other one of the first-mentioned and further hybrid exhaust fluid generators.
In another hypostasis of the invention, the vehicle includes a source of fluid oxidizer, and the control arrangement includes a controllable valve arrangement coupled between the source of fluid oxidizer and the oxidizer input ports of the first-mentioned and further hybrid exhaust fluid generators.
A method for directive control of a vehicle according to another aspect of the invention includes the step of generating high-temperature propulsion fluid, and directing the propulsion fluid through a nozzle to thereby generate thrust along a thrust axis. A fuel grain of a hybrid exhaust gas generator is kept hot by at least one of (a) heating by the high-temperature propulsion fluid, or (b) by a flow of a trickle of oxidizer which combusts with the fuel grain. The fuel grain is thus in a hot state, ready for substantially instantaneous combustion with a flow, or substantial flow greater than the trickle, of oxidizer. When thrust vector modification is desired, substantial oxidizer is supplied to the grain, which combusts, to generate exhaust gas. The exhaust gas is injected or allowed to enter the nozzle in an asymmetrical manner, where it disrupts the flow of the propulsion fluid in a manner which affects the thrust vector.